Rotorcraft control mode transition smoothing

ABSTRACT

A method of operating a rotorcraft includes transitioning from a first mode to a second mode when a velocity of the rotorcraft exceeds a first velocity threshold. Transitioning between the first and second modes includes fading out a gain of a dynamic controller over a first period of time, and decreasing a value of an integrator of the dynamic controller over a second period of time. In the first mode, the translational speed of the rotorcraft is determined based on a pilot stick signal, and in the second mode, an output of an attitude rate controller is proportional to an amplitude of the pilot stick signal.

TECHNICAL FIELD

The present invention relates generally to a system and method for a flight control, and, in particular embodiments, to a system and method for control mode transition smoothing for a rotorcraft.

BACKGROUND

Fly-by-wire systems in aircraft, as opposed to mechanically controlled systems, use electronic signals to control the flight surfaces and engines in the aircraft. For example, instead of having the pilot controls mechanically linked to the control surfaces via a hydraulic system, the pilot controls are electronically linked to a flight computer, which, in turn, controls flight surface actuators via electronic signals. By further interfacing the flight computer to aircraft sensors, sophisticated control algorithms may be used to provide autopilot functionality, as well as to stabilize and control the aircraft.

While fly-by-wire systems have become commonplace in commercial and civilian fixed wing aircraft, their adoption among rotorcraft, such as helicopters, has been much slower due, in part, to the increased complexity of controlling and stabilizing a rotorcraft. However, by adopting fly-by-wire systems in helicopters, safer operation may be achieved in difficult flight environments such as low speed, low altitude, degraded visual environments and inclement weather. Another area in which fly-by-wire systems may benefit rotorcraft is in the reduction in pilot workload. By providing automated features such as stabilization in response to wind, control axis decoupling, position hold and heading hold functionality, the pilot is freed up to focus on the environment in which he flies.

One challenge in the design of fly-by-wire systems for rotorcraft is transitioning between different modes of operation that utilize different control laws or algorithms. In some circumstances, the change in control algorithm may result in a physical transient during operation of the rotorcraft that might be physically discernable as a bump or jolt by the pilot or passengers.

SUMMARY

In accordance with an embodiment, a method of operating a rotorcraft includes transitioning from a first mode to a second mode when a velocity of the rotorcraft exceeds a first velocity threshold. Transitioning between the first and second modes includes fading out a gain of a dynamic controller over a first period of time, and decreasing a value of an integrator of the dynamic controller over a second period of time. In the first mode, the translational speed of the rotorcraft is determined based on a pilot stick signal, and in the second mode, an output of an attitude rate controller is proportional to an amplitude of the pilot stick signal.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the present invention, and the advantages thereof, reference is now made to the following descriptions taken in conjunction with the accompanying drawings, in which:

FIG. 1 illustrates an embodiment rotorcraft;

FIG. 2 illustrates a block diagram of an embodiment rotorcraft flight control system;

FIG. 3 illustrates a block diagram of an embodiment flight control system;

FIG. 4 illustrates a block diagram of a further embodiment flight control system;

FIG. 5 illustrates a block diagram of an embodiment TRC mode attitude controller;

FIG. 6 illustrates a block diagram of an embodiment method; and

FIG. 7 illustrates an embodiment computer system.

Corresponding numerals and symbols in different figures generally refer to corresponding parts unless otherwise indicated. The figures are drawn to clearly illustrate the relevant aspects of the preferred embodiments and are not necessarily drawn to scale. To more clearly illustrate certain embodiments, a letter indicating variations of the same structure, material, or process step may follow a figure number.

DETAILED DESCRIPTION OF ILLUSTRATIVE EMBODIMENTS

Illustrative embodiments of the system and method of the present disclosure are described below. In the interest of clarity, all features of an actual implementation may not be described in this specification. It will of course be appreciated that in the development of any such actual embodiment, numerous implementation-specific decisions may be made to achieve the developer's specific goals, such as compliance with system-related and business-related constraints, which will vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time-consuming but would nevertheless be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure.

Reference may be made herein to the spatial relationships between various components and to the spatial orientation of various aspects of components as the devices are depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of the present disclosure, the devices, members, apparatuses, etc. described herein may be positioned in any desired orientation. Thus, the use of terms such as “above,” “below,” “upper,” “lower,” or other like terms to describe a spatial relationship between various components or to describe the spatial orientation of aspects of such components should be understood to describe a relative relationship between the components or a spatial orientation of aspects of such components, respectively, as the device described herein may be oriented in any desired direction.

The increasing use of rotorcraft, in particular, for commercial and industrial applications, has led to the development of larger more complex rotorcraft. However, as rotorcraft become larger and more complex, the differences between flying rotorcraft and fixed wing aircraft has become more pronounced. Since rotorcraft use one or more main rotors to simultaneously provide lift, control attitude, control altitude, and provide lateral or positional movement, different flight parameters and controls are tightly coupled to each other, as the aerodynamic characteristics of the main rotors affect each control and movement axis. For example, the flight characteristics of a rotorcraft at cruising speed or high speed may be significantly different than the flight characteristics at hover or at relatively low speeds. Additionally, different flight control inputs for different axes on the main rotor, such as cyclic inputs or collective inputs, affect other flight controls or flight characteristics of the rotorcraft. For example, pitching the nose of a rotorcraft forward to increase forward speed will generally cause the rotorcraft to lose altitude. In such a situation, the collective may be increased to maintain level flight, but the increase in collective causes increased power to the main rotor which, in turn, requires additional anti-torque force from the tail rotor. This is in contrast to fixed wing systems where the control inputs are less closely tied to each other and flight characteristics in different speed regimes are more closely related to each other.

Recently, fly-by-wire (FBW) systems have been introduced in rotorcraft to assist pilots in stably flying the rotorcraft and to reduce workload on the pilots. The FBW system may provide different control characteristics or responses for cyclic, pedal or collective control input in the different flight regimes, and may provide stability assistance or enhancement by decoupling physical flight characteristics so that a pilot is relieved from needing to compensate for some flight commands issued to the rotorcraft. FBW systems may be implemented in one or more flight control computers (FCCs) disposed between the pilot controls and flight control systems, providing corrections to flight controls that assist in operating the rotorcraft more efficiently or that put the rotorcraft into a stable flight mode while still allowing the pilot to override the FBW control inputs. The FBW systems in a rotorcraft may, for example, automatically adjust power output by the engine to match a collective control input, apply collective or power correction during a cyclic control input, provide automation of one or more flight control procedures, provide for default or suggested control positioning, or the like.

FBW systems for rotorcraft must provide stable flight characteristics for FBW controlled flight parameters while permitting the pilot to override or adjust any suggested flight parameters suggested by the FBW system. Additionally, in providing enhanced control and automated functionality for rotorcraft flight, the FBW system must maintain an intuitive and easy to use flight control system for the pilot. Thus, the FBW system adjusts the pilot flight controls so that the controls are in a position associated with the relevant flight parameter. For example, the FBW system may adjust the collective stick to provide suggested or FBW controlled flight parameters, and which reflect a collective or power setting. Thus, when the pilot releases the collective stick and the FBW system provides collective control commands, the collective stick is positioned intuitively in relation to the actual power or collective setting so that, when the pilot grasps the collective stick to retake control, the control stick is positioned where the pilot expects the stick to be positioned for the actual collective setting of the main rotor. Similarly, the FBW system use the cyclic stick to, for example, adjust for turbulence, drift or other disturbance to the flight path, and may move the cyclic stick as the FBW system compensates the cyclic control. Thus, when the pilot grasps the cyclic stick to take control of flight from the FBW system, the cyclic stick is positioned to reflect the actual cyclic settings.

Embodiments of the present disclosure will be described with respect to preferred embodiments in a specific context, namely a system and method for smoothing a transition between a translational rate command (TRC) mode to a rate mode in a rotorcraft. Embodiments of the present disclosure may also be applied to other control mode transitions in the operation and control of a rotorcraft.

In an embodiment of the present invention, a forward displacement on the cyclic controller or pilot stick of the rotorcraft is interpreted in one of two ways. When the speed of the rotorcraft is 10 kts or less and the speed command by displacement on the cyclic controller is less than 10 kts, the displacement on the cyclic controller is interpreted as a speed command. For example, ½ inch of stick displacement may correspond to a speed of 2 kts; 1″ of stick displacement may correspond to a speed of 4 kts, and so on. However, when the speed of the rotorcraft is greater than 10 kts or the speed command by displacement on the cyclic controller is greater than 10 kts, displacement on the cyclic controller is interpreted as a rate command. For example, ½″ of stick displacement may correspond to a rate of 10° per second; 1″ of stick displacement may correspond to a rate of 20° per second, and so on. By automatically switching between the two different modes of operation based on the speed of the rotorcraft, pilot workload may be reduced.

In some circumstances, however, the rotorcraft may undergo a physical transient during the transition between the speed control mode at low speeds and the rate control mode at higher speeds. This physical transient may be caused by sudden changes in flight control algorithms and abrupt shifts in control signals. In embodiments of the present invention, the transition between speed control and rate control is made smoother by controlling the gains and the signal path structure of an attitude controller used during the speed control mode. In one example, the overall gain of the attitude controller is reduced using a fader. During this time, a proportional path in the controller is maintained while a value of an integrator is decremented to zero. Also during this time the gain of a control path between the cyclic controller and the rate control block is increased using a fader. By maintaining a proportional path in the attitude controller, decrementing the integrator to zero, fading out the overall gain of the attitude controller and fading in the gain of the control path between the cyclic controller and the rate control block, a smooth transition between speed control mode and rate control mode may be made more smooth. It should be understood that embodiments of the present invention may be applied to other control signals and control paths of the rotorcraft.

FIG. 1 illustrates a rotorcraft 101 according to some embodiments. The rotorcraft 101 has a main rotor system 103, which includes a plurality of main rotor blades 105. The pitch of each main rotor blade 105 may be controlled by a swashplate 107 in order to selectively control the attitude, altitude and movement of the rotorcraft 101. The swashplate 107 may be used to collectively and/or cyclically change the pitch of the main rotor blades 105. The rotorcraft 101 also has an anti-torque system, which may include a tail rotor 109, no-tail-rotor (NOTAR), or dual main rotor system. In rotorcraft with a tail rotor 109, the pitch of each tail rotor blade 111 is collectively changed in order to vary thrust of the anti-torque system, providing directional control of the rotorcraft 101. The pitch of the tail rotor blades 111 is changed by one or more tail rotor actuators. In some embodiments, the FBW system sends electrical signals to the tail rotor actuators or main rotor actuators to control flight of the rotorcraft.

Power is supplied to the main rotor system 103 and the anti-torque system by engines 115. There may be one or more engines 115, which may be controlled according to signals from the FBW system. The output of the engine 115 is provided to a driveshaft 117, which is mechanically and operatively coupled to the rotor system 103 and the anti-torque system through a main rotor transmission 119 and a tail rotor transmission, respectively.

The rotorcraft 101 further includes a fuselage 125 and tail section 123. The tail section 123 may have other flight control devices such as horizontal or vertical stabilizers, rudder, elevators, or other control or stabilizing surfaces that are used to control or stabilize flight of the rotorcraft 101. The fuselage 125 includes a cockpit 127, which includes displays, controls, and instruments. It should be appreciated that even though rotorcraft 101 is depicted as having certain illustrated features, the rotorcraft 101 may have a variety of implementation-specific configurations. For instance, in some embodiments, cockpit 127 is configured to accommodate a pilot or a pilot and co-pilot, as illustrated. It is also contemplated, however, that rotorcraft 101 may be operated remotely, in which case cockpit 127 could be configured as a fully functioning cockpit to accommodate a pilot (and possibly a co-pilot as well) to provide for greater flexibility of use, or could be configured with a cockpit having limited functionality (e.g., a cockpit with accommodations for only one person who would function as the pilot operating perhaps with a remote co-pilot or who would function as a co-pilot or back-up pilot with the primary piloting functions being performed remotely). In yet other contemplated embodiments, rotorcraft 101 could be configured as an unmanned vehicle, in which case cockpit 127 could be eliminated entirely in order to save space and cost.

FIG. 2 illustrates a fly-by-wire flight control system 201 for a rotorcraft according to some embodiments. A pilot may manipulate one or more pilot flight controls in order to control flight of the rotorcraft. The pilot flight controls may include manual controls such as a cyclic stick 231 in a cyclic control assembly 217, a collective stick 233 in a collective control assembly 219, and pedals 239 in a pedal assembly 221. Inputs provided by the pilot to the pilot flight controls may be transmitted mechanically and/or electronically (e.g., via the FBW flight control system) to flight control devices by the flight control system 201. Flight control devices may represent devices operable to change the flight characteristics of the rotorcraft. Flight control devices on the rotorcraft may include mechanical and/or electrical systems operable to change the positions or angle of attack of the main rotor blades 105 and the tail rotor blades in or to change the power output of the engines 115, as examples. Flight control devices include systems such as the swashplate 107, tail rotor actuator 113, and systems operable to control the engines 115. The flight control system 201 may adjust the flight control devices independently of the flight crew in order to stabilize the rotorcraft, reduce workload of the flight crew, and the like. The flight control system 201 includes engine control computers (ECCUs) 203, flight control computers 205, and aircraft sensors 207, which collectively adjust the flight control devices.

The flight control system 201 has one or more flight control computers 205 (FCCs). In some embodiments, multiple FCCs 205 are provided for redundancy. One or more modules within the FCCs 205 may be partially or wholly embodied as software and/or hardware for performing any functionality described herein. In embodiments where the flight control system 201 is a FBW flight control system, the FCCs 205 may analyze pilot inputs and dispatch corresponding commands to the ECCUs 203, the tail rotor actuator 113, and/or actuators for the swashplate 107. Further, the FCCs 205 are configured and receive input commands from the pilot controls through sensors associated with each of the pilot flight controls. The input commands are received by measuring the positions of the pilot controls. The FCCs 205 also control tactile cueing commands to the pilot controls or display information in instruments on, for example, an instrument panel 241.

The ECCUs 203 control the engines 115. For example, the ECCUs 203 may vary the output power of the engines 115 to control the rotational speed of the main rotor blades or the tail rotor blades. The ECCUs 203 may control the output power of the engines 115 according to commands from the FCCs 205, or may do so based on feedback such a measured revolutions per minute (RPM) of the main rotor blades.

The aircraft sensors 207 are in communication with the FCCs 205. The aircraft sensors 207 may include sensors for measuring a variety of rotorcraft systems, flight parameters, environmental conditions and the like. For example, the aircraft sensors 207 may include sensors for measuring airspeed, altitude, attitude, position, orientation, temperature, airspeed, vertical speed, and the like. Other sensors 207 could include sensors relying upon data or signals originating external to the rotorcraft, such as a global positioning system (GPS) sensor, a VHF Omnidirectional Range sensor, Instrument Landing System (ILS), and the like.

The cyclic control assembly 217 is connected to a cyclic trim assembly 229 having one or more cyclic position sensors 211, one or more cyclic detent sensors 235, and one or more cyclic actuators or cyclic trim motors 209. The cyclic position sensors 211 measure the position of the cyclic control stick 231. In some embodiments, the cyclic control stick 231 is a single control stick that moves along two axes and permits a pilot to control pitch, which is the vertical angle of the nose of the rotorcraft and roll, which is the side-to-side angle of the rotorcraft. In some embodiments, the cyclic control assembly 217 has separate cyclic position sensors 211 that measuring roll and pitch separately. The cyclic position sensors 211 for detecting roll and pitch generate roll and pitch signals, respectively, (sometimes referred to as cyclic longitude and cyclic latitude signals, respectively) which are sent to the FCCs 205, which controls the swashplate 107, engines 115, tail rotor 109 or related flight control devices.

The cyclic trim motors 209 are connected to the FCCs 205, and receive signals from the FCCs 205 to move the cyclic control stick 231. In some embodiments, the FCCs 205 determine a suggested cyclic stick position for the cyclic stick 231 according to one or more of the collective stick position, the pedal position, the speed, altitude and attitude of the rotorcraft, the engine revolutions per minute (RPM), engine temperature, main rotor RPM, engine torque or other rotorcraft system conditions or flight conditions. The suggested cyclic stick position is a positon determined by the FCCs 205 to give a desired cyclic action. In some embodiments, the FCCs 205 send a suggested cyclic stick position signal indicating the suggested cyclic stick position to the cyclic trim motors 209. While the FCCs 205 may command the cyclic trim motors 209 to move the cyclic stick 231 to a particular position (which would in turn drive actuators associated with swashplate 107 accordingly), the cyclic position sensors 211 detect the actual position of the cyclic stick 231 that is set by the cyclic trim motors 206 or input by the pilot, allowing the pilot to override the suggested cyclic stick position. The cyclic trim motor 209 is connected to the cyclic stick 231 so that the pilot may move the cyclic stick 231 while the trim motor is driving the cyclic stick 231 to override the suggested cyclic stick position. Thus, in some embodiments, the FCCs 205 receive a signal from the cyclic position sensors 211 indicating the actual cyclic stick position, and do not rely on the suggested cyclic stick position to command the swashplate 107.

Similar to the cyclic control assembly 217, the collective control assembly 219 is connected to a collective trim assembly 225 having one or more collective position sensors 215, one or more collective detent sensors 237, and one or more collective actuators or collective trim motors 213. The collective position sensors 215 measure the position of a collective control stick 233 in the collective control assembly 219. In some embodiments, the collective control stick 233 is a single control stick that moves along a single axis or with a lever type action. A collective position sensor 215 detects the position of the collective control stick 233 and sends a collective position signal to the FCCs 205, which controls engines 115, swashplate actuators, or related flight control devices according to the collective position signal to control the vertical movement of the rotorcraft. In some embodiments, the FCCs 205 may send a power command signal to the ECCUs 203 and a collective command signal to the main rotor or swashplate actuators so that the angle of attack of the main blades is raised or lowered collectively, and the engine power is set to provide the needed power to keep the main rotor RPM substantially constant.

The collective trim motor 213 is connected to the FCCs 205, and receives signals from the FCCs 205 to move the collective control stick 233. Similar to the determination of the suggested cyclic stick position, in some embodiments, the FCCs 205 determine a suggested collective stick position for the collective control stick 233 according to one or more of the cyclic stick position, the pedal position, the speed, altitude and attitude of the rotorcraft, the engine RPM, engine temperature, main rotor RPM, engine torque or other rotorcraft system conditions or flight conditions. The FCCs 205 generate the suggested collective stick position and send a corresponding suggested collective stick signal to the collective trim motors 213 to move the collective stick 233 to a particular position. The collective position sensors 215 detect the actual position of the collective stick 233 that is set by the collective trim motor 213 or input by the pilot, allowing the pilot to override the suggested collective stick position.

The pedal control assembly 221 has one or more pedal sensors 227 that measure the position of pedals or other input elements in the pedal control assembly 221. In some embodiments, the pedal control assembly 221 is free of a trim motor or actuator, and may have a mechanical return element that centers the pedals when the pilot releases the pedals. In other embodiments, the pedal control assembly 221 has one or more trim motors that drive the pedal to a suggested pedal position according to a signal from the FCCs 205. The pedal sensor 227 detects the position of the pedals 239 and sends a pedal position signal to the FCCs 205, which controls the tail rotor 109 to cause the rotorcraft to yaw or rotate around a vertical axis.

The cyclic and collective trim motors 209 and 213 may drive the cyclic stick 231 and collective stick 233, respectively, to suggested positions. The cyclic and collective trim motors 209 and 213 may drive the cyclic stick 231 and collective stick 233, respectively, to suggested positions, but this movement capability may also be used to provide tactile cueing to a pilot. The trim motors 209 and 213 may push the respective stick in a particular direction when the pilot is moving the stick to indicate a particular condition. Since the FBW system mechanically disconnects the stick from one or more flight control devices, a pilot may not feel a hard stop, vibration, or other tactile cue that would be inherent in a stick that is mechanically connected to a flight control assembly. In some embodiments, the FCCs 205 may cause the trim motors 209 and 213 to push against a pilot command so that the pilot feels a resistive force, or may command one or more friction devices to provide friction that is felt when the pilot moves the stick. Thus, the FCCs 205 control the feel of a stick by providing pressure and/or friction on the stick.

Additionally, the cyclic control assembly 217, collective control assembly 219 and/or pedal control assembly 221 may each have one or more detent sensors that determine whether the pilot is handling a particular control device. For example, the cyclic control assembly 217 may have a cyclic detent sensor 235 that determines that the pilot is either holding or moving the cyclic stick 231, while the collective control assembly 219 has a collective detent sensor 237 that determines whether the pilot is either holding or moving the collective stick 233. These detent sensors 235, 237 detect motion and/or position of the respective control stick that is caused by pilot input, as opposed to motion and/or position caused by commands from the FCCs 205, rotorcraft vibration, and the like and provide feedback signals indicative of such to the FCCs. When the FCCs 205 detect that a pilot has control of, or is manipulating, a particular control, the FCCs 205 may determine that stick to be out-of-detent (00D). Likewise, the FCCs may determine that the stick is in-detent (ID) when the signals from the detent sensors indicate to the FCCs 205 that the pilot has released a particular stick. The FCCs 205 may provide different default control or automated commands to one or more flight systems based on the detent status of a particular stick or pilot control.

Moving now to the operational aspects of flight control system 201, FIG. 3 illustrates in a highly schematic fashion, a manner in which flight control system 201 may implement FBW functions as a series of inter-related feedback loops running certain control laws. FIG. 3 representatively illustrates a three-loop flight control system 201 according to an embodiment. In some embodiments, elements of the three-loop flight control system 201 may be implemented at least partially by FCCs 205. As shown in FIG. 3, however, all, some, or none of the components (301, 303, 305, 307) of three-loop flight control system 201 could be located external or remote from the rotorcraft 100 and communicate to on-board devices through a network connection 309.

The three-loop flight control system 201 of FIG. 3 has a pilot input 311, an outer loop 313, a rate (middle) loop 315, an inner loop 317, a decoupler 319, and aircraft equipment 321 (corresponding, e.g., to flight control devices such as swashplate 107, tail rotor transmission 212, etc., to actuators (not shown) driving the flight control devices, to sensors such as aircraft sensors 207, position sensors 211, 215, detent sensors 235, 237, etc., and the like).

In the example of FIG. 3, a three-loop design separates the inner stabilization and rate feedback loops from outer guidance and tracking loops. The control law structure primarily assigns the overall stabilization task and related tasks of reducing pilot workload to inner loop 317. Next, middle loop 315 provides rate augmentation. Outer loop 313 focuses on guidance and tracking tasks. Since inner loop 317 and rate loop 315 provide most of the stabilization, less control effort is required at the outer loop level. As representatively illustrated in FIG. 3, a switch 322 may be provided to turn outer loop flight augmentation on and off, as the tasks of outer loop 313 are not necessary for flight stabilization.

In some embodiments, the inner loop 317 and rate loop 315 include a set of gains and filters applied to roll/pitch/yaw 3-axis rate gyro and acceleration feedback sensors. Both the inner loop and rate loop may stay active, independent of various outer loop hold modes. Outer loop 313 may include cascaded layers of loops, including an attitude loop, a speed loop, a position loop, a vertical speed loop, an altitude loop, and a heading loop. In accordance with some embodiments, the control laws running in the illustrated the loops allow for decoupling of otherwise coupled flight characteristics, which in turn may provide for more stable flight characteristics and reduced pilot workload. Furthermore, the outer loop 313 may allow for automated or semi-automated operation of certain high-level tasks or flight patterns, thus further relieving the pilot workload and allowing the pilot to focus on other matters including observation of the surrounding terrain.

FIG. 4 illustrates a flight control system 400 according to an embodiment of the present invention. Pilot stick pitch block 401 a and pilot stick roll block 401 b represent, for example, the respective pitch and roll commands emanating from cyclic controller 217 of the rotorcraft. As shown, pilot stick blocks 401 a and 401 b interface to flight controller 402. In various embodiments, flight controller 402 is implemented using a flight computer, or other processing hardware. Flight controller 402 also interfaces with and controls aircraft equipment 321 representing various actuators, sensors, and the physical body of the rotorcraft. In various embodiments, flight controller 402 controls aircraft equipment 321 using three loops: an inner loop, a rate feedback loop and a state feedback loop: the inner loop stabilizes the dynamics of the rotorcraft, the rate loop controls the angular rates of the rotor craft, and the outer loop provides control signals to the inner loop and/or rate loops to effect a desired attitude, speed and position of the rotorcraft. In some embodiments, the outer loop supports and provides flight augmentation or auto-pilot functionality and may be manually or automatically disabled based on flight and system conditions. The inner loop and rate feedback loops, on the other hand, remain operational to provide stability to the rotorcraft.

For purposes of illustration, flight controller 402 is illustrated with respect to the general control blocks that affect the rotational rate of an embodiment rotorcraft, namely the control blocks affecting the roll rate and the pitch rate of the rotorcraft. It should be understood that flight controller 402 may also include other controllers and control paths that affect the yaw and other states of the rotorcraft in addition to the pitch rate and roll rate. As shown, the inner stabilization loop is controlled by inner loop controller 317, the rate loop is controlled by pitch rate controller 315 a and roll rate controller 315 b, and the outer loop is controlled by outer loop controller 313 in conjunction with TRC mode attitude controllers 408 a and 408 b that provide TRC mode pitch and roll control, respectively, during embodiment TRC modes.

Each of inner loop controller 317, decoupler 319 and rate controller 315 may be implemented using flight control algorithms known in the art. Inner loop controller 317 receives sensor feedback from sensors such as gyroscopes and accelerometers within the rotorcraft and provides control signals to various actuators, such as swashplate 107 to stabilize the rotorcraft. Rate controller 315 receives rate feedback from rate gyroscopes on all axes and provides a rate command signal to inner loop controller 317 based on the rate feedback and the position of pilot stick pitch block 401 a and pilot stick roll block 401 b in some modes of operation. Decoupler 319 receives the various rate commands and approximately decouples all 4-axes (pitch, roll, yaw, and vertical) such that, for example, a forward longitudinal stick input does not require the pilot to push the stick diagonally.

Outer loop controller 313 receives state feedback from the sensors of aircraft equipment 321. This state feedback may include, for example, speed, position and attitude. In a translational rate commend (TRC) mode, outer loop controller 313 receives a command from the pilot stick represented by pilot stick pitch block 401 a and pilot stick roll block 401 b, determines a corresponding translational speed based on the pilot stick command, and determines corresponding pitch and roll attitude commands based on the determined translational speed. The translational speed may include a forward component based on commands from pilot stick pitch block 401 a and a lateral component based on commands from pilot stick roll block 401 b. In some embodiments, the corresponding translational speed is determined by multiplying the pilot stick command, or a vector sum of the pitch and roll pilot stick commands with a scale factor and/or by using a lookup table. The translational speed may be calculated using a vector sum as follows: V _(T) =k√{square root over (S _(LON) ² +S _(LAT) ²)}, where V_(T) is the translational speed, k is a scale factor, S_(LON) is the pilot stick pitch command, and S_(LAT) is the pilot stick roll command. In some embodiments, the vector sum is not used and the translational speed is based on the forward speed and/or the lateral speed.

The pitch and roll attitude commands may be determined from the translational speed using a speed control loop. During the TRC mode, TRC mode attitude controllers 408 a and 408 b individually calculate pitch and roll attitude errors by subtracting the pitch attitude feedback from the pitch attitude command and by subtracting the roll attitude feedback from the roll attitude command. In various embodiments, the pitch and roll attitude feedback is a component of the state feedback. TRC mode attitude controller 408 a applies a dynamic control algorithm to the pitch attitude error to produce an outer loop pitch rate command, and TRC mode attitude controller 408 b applies a dynamic control algorithm to the roll attitude error to produce an outer loop roll rate command. These outer loop rate commands are applied to decoupler 319 via summing blocks 412 a and 412 b. While the embodiment of FIG. 4 shows TRC mode attitude controllers 408 a and 408 b used to provide TRC mode control to both the pitch and roll channels, it should be understood that in alternative embodiments, TRC mode control may be applied to a single attitude channel, such as the pitch channel for speed control in the forward direction.

In various embodiments of the present invention, the TRC mode is automatically selected based on the velocity or ground speed of the rotorcraft. This ground speed may be measured, for example, using GPS. In one example, the TRC mode is selected when the ground speed is less than 10 kts, in which a fixed offset of the pilot stick effects a constant translation. When the ground speed is 10 kts and greater, the pilot stick effects a rate in a rate control mode. In the embodiment of FIG. 4, the TRC mode is activated by enabling TRC mode attitude controller 408 a and 408 b when the ground speed is less than 10 kts and decoupling the pilot stick pitch block 401 a from pitch rate controller 315 a via fader 404 a and decoupling the pilot stick roll block 401 b from the roll rate controller 315 b via fader 404 b. It should be appreciated that the 10 kts threshold is just one example of many possible groundspeed thresholds. In alternative embodiments, other threshold values may be used. Likewise, the rate control mode is selected by coupling pilot stick pitch block 401 a to pitch rate controller 315 a via fader 404 a, and by coupling pilot stick roll block 401 b to roll rate controller 315 a to pitch rate controller via fader 404 b. In some embodiments, pilot stick blocks 401 a and 401 b are coupled to rate controllers 315 a and 315 b when the pilot stick is out of detent. In some embodiments, hysteresis may be applied to the ground speed threshold to prevent metastability of the control mode when the rotorcraft is flying at about 10 kts. In such an embodiment the ground speed threshold for transitioning from the TRC to the rate control mode is greater than the ground speed threshold for transitioning from the rate control mode to the TRC mode. For example, in one embodiment, TRC mode attitude controllers 408 a and 408 b are disabled when the pilot stick command is greater than 10 kts and then re-enabled when the pilot stick command is less than 3 kts. Alternatively, other hysteresis/threshold values may be used. Alternatively, a single threshold may be used without hysteresis. In various embodiments, the 10 kts and 3 kts speed thresholds may be applied to the forward speed, the lateral speed or a combination of the forward and lateral speed. Such a combination may be determined, for example, by calculating a vector sum of the forward speed and the lateral speed.

During operation, faders 404 a and 404 b provide a gain of one when the pilot stick is out of detent, and provides a gain of zero when the pilot stick is in detent. In various embodiments, the gain of faders 404 a and 404 b linearly increase or decrease between zero and one over a predetermined period of time. This predetermined period of time is between about 1 s and about 10 s; however, fading times outside of this range may be implemented depending on the particular embodiment and its specifications. In further alternative embodiments, the gains of fader 404 a and 404 b may be different from one and zero. In some embodiments, pitch rate controllers 315 a and roll rate controller 315 b each a high-pass filter and/or washout filter coupled to their respective inputs. Thus, when the stick is out of detent and held at a steady displacement, the effective command produced by rate controller 315 a and 315 b as a result of a pilot stick command goes to zero. This high-pass filter or washout filter also prevents the output of TRC mode attitude controller 408 a and 408 b and the output of faders 404 a and 404 b from conflicting or “fighting” with each other when the pilot stick is out of detent.

FIG. 5 illustrates a block diagram of TRC mode attitude controller 408 according to an embodiment of the present invention that may be used to implement TRC attitude controllers 408 a and 408 b illustrated in FIG. 4. As shown, TRC mode attitude controller 408 includes subtractor 430, fader 432, dynamic compensation block 434 and proportional-integral (PI) controller 450. Subtractor 430 produces an error signal based on the attitude command produced by the outer loop controller 313 (FIG. 4) and dynamic compensation block 434 compensates for the dynamics of the rotorcraft in order to improve stability and/or adjust the time response of the attitude loop. Dynamic compensation block 434 may include various control blocks known in the art including but not limited a PID controller and a lead-lag compensator.

PI controller 450 includes a proportional signal path that includes proportional gain block 436 coupled in parallel with an integral signal path that includes integral gain block 438 and integrator 440. In various embodiments, proportional gain block 436 and integral gain block 438 may be implemented, for example, by performing a multiplication or scaling operation. Integrator 440 may be implemented, for example, using an accumulator.

During TRC mode operation below the 10 kts speed threshold, fader 432 has a gain of one and the input of integrator 440 is coupled to the output of dynamic compensation block 434 via integral gain block 438. On the other hand, in the rate control mode in which the rotorcraft has a ground speed of 10 kts and above or the speed command by displacement on the cyclic controller is greater than 10 kts, fader 432 has a gain of zero, and the input of integrator 440 is coupled to its output via feedback gain block 444 and limiter 442, which effectively causes integrator 440 to decrement to zero or to a DC value representing zero output. The combination of zero fader gain and zero integrator output effectively disables TRC mode attitude controller 408. When the ground speed of the rotorcraft exceeds the 10 kts threshold, the gain of fader 432 linearly decreases to zero over a predetermined time period and the input of the integrator is switched from the output of integral gain block 438 to the output of limiter 442. Once the input to integrator is switched to the output of limiter 442 via switch 452, the feedback action of the loop formed by integrator 440, feedback gain 444 and limiter 442 forces the output of integrator 440 decrement to zero over a period of time. Limiter 442 limits the rate at which integrator 440 is decremented such that the decay of integrator 440 has more of a ramp response instead of an exponential response for high integrator output values. Thus, in some embodiments, during a mode transition from TRC mode to rate control mode, the actuator command generated by summer 446 produces a gently decreasing or decaying signal from the forward path of the TRC mode attitude controller 408 due to fader 432 that is summed with a gently decreasing or decaying integrator value produced by integrator 440. At the same time as the output of the TRC mode attitude controller 408 is decaying, the pilot stick command to rate controller 315 is increasing due the increasing gain of fader 404. The net result of the decreasing output of the TRC mode attitude controller 408 and the increasing output of fader 404 a or 404 b to rate controller 315 is a smooth transition from TRC mode to attitude mode in some embodiments. In various embodiments, the time that it takes for faders 404 a, 404 b and 432 to change their gains may be the same or may be different from each other. In some embodiments, the fade-in and fade-out times of faders 404 a, 404 b and 432 may be the same or they may be different from each other.

Similarly, during a mode transition from rate mode to TRC mode, the pilot stick control signal produced by fader 404 a or 404 b decreases while the forward path of TRC mode attitude controller 408 increases due to the gain of fader 432 increasing and due to the integral path of PI controller 450 being activated by coupling the input of integrator 440 to the output of integral gain block 438 via switch 452.

In alternative embodiments of the present invention, integrator may be decremented in a different manner. For example, instead of using feedback gain 444 to decrease the output of the integrator, a fader may be coupled to the output of integrator 440 or may be coupled after summer 446.

In various embodiments, the transfer function and implementation of dynamic compensation block 434, the gains of fader 432, proportional gain 435, integral gain 438 and feedback gain 444, and the static and dynamic behavior of fader 432, limiter 442 and integrator 440 may be adjusted according to the specification and requirements of the rotorcraft and may have different implementations and values with regard to the roll, pitch and other attitude channels.

It should be understood that the transition between TRC and rate control mode with respect to the various rate commands is just one example of many possible system configurations that utilize embodiment mode switching systems and methods. In alternative embodiments, other control channels besides the pitch rate and roll rate control channel may be controlled. For example, embodiments of the present invention may be used to transition between attitude hold/command mode and a speed hold/command mode, transition between attitude hold/command mode and a position hold/command mode, transition between direct vertical collective control mode and an altitude hold/command mode, and/or transition between vertical speed hold/command mode and an altitude hold/command mode. It should be appreciated that these are only a few examples of many possible applications.

FIG. 6 illustrates a block diagram of an embodiment method 500 of operating a rotorcraft that may be executed, for example, by a flight computer of a fly-by-wire system. In an embodiment, the pilot stick command of the cyclic control is interpreted as a translational rate command in step 518 in which the flight computer controls the rotorcraft to have a translational velocity that is proportional to a physical offset of the collective control. During step 518, a dynamic controller adjusts the attitude of the rotorcraft in order to maintain the desired translational velocity. When the velocity of the rotorcraft exceeds a threshold V_(threshold2) according to step 520 or when the pilot stick command exceeds V_(threshold2), the operational mode of the flight computer transitions to a rate command mode in step 508 in which the pilot stick command of the cyclic control is interpreted as a rate command. In this mode, the flight computer controls the rotorcraft to have a rate that is proportional to the physical offset of the collective control.

In order to provide a smooth transition from the translational command mode of step 518 to the rate command mode of step 508 and to reduce the occurrence of physical transients of the rotorcraft, steps 502, 504 and 506 are executed by the flight computer. In step 502, the gain of a first dynamic controller, which may be implemented, for example, as outer loop controller that provides a rate control in response to a desired attitude command, is faded out over first period of time. In step 504, a state of an integrator of the first dynamic controller is decreases over a second period of time, and in step 506 a gain of the pilot stick command is faded in over a third period of time. In some embodiments, steps 502, 504 and 506 occur concurrently such that the gain of the first dynamic controller is reduced at the same time as the gain of the pilot stick command to the rate controller is faded in. During this time, the value of the integrator is reduced. In some embodiments, the value of the integrator is decremented by disconnecting the integrator from the input of the first dynamic controller and coupling the input of the integrator to its output. The input and/or output value of the integrator may be limited while the integrator value is reduced in order to increase the amount of time that it takes to decrement the value of the integrator to zero.

The rate command continues to be determined by the pilot stick command in step 508 until the velocity of the rotorcraft is less than or equal to threshold V_(threshold1) and the stick command S_(command) is less than an offset S_(threshold) that represents a particular translational velocity command according to step 510. In some embodiments, S_(threshold) may be set to a pilot stick offset that is less than 3 kts. Alternatively, other thresholds may be used. Once this condition is detected, the operational mode of the flight computer transitions to the translational rate command mode of step 518 via steps 512, 514 and 516. In some embodiments, threshold V_(threshold1) is equal to threshold V_(threshold2), while in other embodiments, V_(threshold2) is greater threshold V_(threshold1) in order to provide hysteresis. In steps 512, the gain of the first dynamic controller is faded in, in step 514, the integrator of the first dynamic controller is reactivated by coupling its input to the signal path of the first dynamic controller, and in step 516 the gain of the pilot stick command to the rate controller is faded out. In some embodiments, steps 512, 514 and 516 occur concurrently such that the gain of the first dynamic controller is increased at the same time as the gain of the pilot stick command to the rate controller is reduced and at the same time the integrator of the first dynamic controller is reactivated.

FIG. 7 illustrates a computer system 601. The computer system 601 can be configured for performing one or more functions with regard to the operation of the flight control system 201 and the method 500, as described herein. Further, any processing and analysis can be partly or fully performed by the computer system 601. The computer system 601 can be partly or fully integrated with other aircraft computer systems or can be partly or fully removed from the rotorcraft.

The computer system 601 can include an input/output (I/O) interface 603, an analysis engine 605, and a database 607. Alternative embodiments can combine or distribute the I/O interface 603, the analysis engine 605, and the database 607, as desired. Embodiments of the computer system 601 may include one or more computers that include one or more processors and memories configured for performing tasks described herein. This can include, for example, a computer having a central processing unit (CPU) and non-volatile memory that stores software instructions for instructing the CPU to perform at least some of the tasks described herein. This can also include, for example, two or more computers that are in communication via a computer network, where one or more of the computers include a CPU and non-volatile memory, and one or more of the computer's non-volatile memory stores software instructions for instructing any of the CPU(s) to perform any of the tasks described herein. Thus, while the exemplary embodiment is described in terms of a discrete machine, it should be appreciated that this description is non-limiting, and that the present description applies equally to numerous other arrangements involving one or more machines performing tasks distributed in any way among the one or more machines. It should also be appreciated that such machines need not be dedicated to performing tasks described herein, but instead can be multi-purpose machines, for example computer workstations, that are suitable for also performing other tasks.

The I/O interface 603 can provide a communication link between external users, systems, and data sources and components of the computer system 601. The I/O interface 603 can be configured for allowing one or more users to input information to the computer system 601 via any known input device. Examples can include a keyboard, mouse, touch screen, and/or any other desired input device. The I/O interface 603 can be configured for allowing one or more users to receive information output from the computer system 601 via any known output device. Examples can include a display monitor, a printer, cockpit display, and/or any other desired output device. The I/O interface 603 can be configured for allowing other systems to communicate with the computer system 601. For example, the I/O interface 603 can allow one or more remote computer(s) to access information, input information, and/or remotely instruct the computer system 601 to perform one or more of the tasks described herein. The I/O interface 603 can be configured for allowing communication with one or more remote data sources. For example, the I/O interface 603 can allow one or more remote data source(s) to access information, input information, and/or remotely instruct the computer system 601 to perform one or more of the tasks described herein.

The database 607 provides persistent data storage for the computer system 601. Although the term “database” is primarily used, a memory or other suitable data storage arrangement may provide the functionality of the database 607. In alternative embodiments, the database 607 can be integral to or separate from the computer system 601 and can operate on one or more computers. The database 607 preferably provides non-volatile data storage for any information suitable to support the operation of the flight control system 201 and the method 500, including various types of data discussed further herein. The analysis engine 605 can include various combinations of one or more processors, memories, and software components.

Embodiments of the present invention are summarized here. Other embodiments can also be understood from the entirety of the specification and the claims filed herein. One general aspect includes a method of operating a rotorcraft that includes operating the rotorcraft in a first mode including determining a translational speed based on a pilot stick signal generated by a pilot stick assembly, determining an attitude based on the determined translational speed, determining an actuator command based on the determined attitude by using a first dynamic controller having an integrator, and providing an output of the first dynamic controller to an actuator, where a translational speed of the rotorcraft is proportional to an amplitude of the pilot stick signal in the first mode. The method also includes transitioning from the first mode to a second mode when a velocity of the rotorcraft exceeds a first velocity threshold, where transitioning includes fading out a gain of the first dynamic controller over a first period of time and decreasing a value of the integrator over a second period of time. The method also includes operating the rotorcraft in the second mode including providing an output of an attitude rate controller to the actuator, where the output of the attitude rate controller is proportional to an amplitude of the pilot stick signal.

Implementations may include one or more of the following features. The method where: determining the translational speed based on a pilot stick signal includes determining a forward speed component based on cyclic longitude stick control portion of the pilot stick signal and determining a lateral speed component based on a cyclic latitude stick control portion of the pilot stick signal; determining the attitude based on the determined translational speed includes determining a pitch attitude based on the determined forward speed component and a roll attitude based on the determined lateral speed component; determining the actuator command includes determining a pitch actuator command based on the determined pitch attitude and a roll actuator command based on the determined roll attitude; the first dynamic controller includes a first pitch dynamic controller having a pitch integrator, and a first roll dynamic controller having a roll integrator; providing the output of the first dynamic controller to the actuator includes providing an output of the first pitch dynamic controller to a pitch actuator and an output of the first roll dynamic controller to a roll actuator; fading out the gain of the first dynamic controller over a first period of time includes fading out a gain of the first pitch dynamic controller and the first roll dynamic controller over the first period of time; and decreasing the value of the integrator over the second period of time includes decreasing a value of the pitch integrator and the roll integrator over the second period of time.

In an embodiment, the translational speed further includes determining a vector sum of the forward speed component and the lateral speed component. In some embodiments, the translational speed includes a forward speed, the attitude includes a pitch attitude, the attitude rate controller includes a pitch rate controller and the actuator includes a pitch actuator. The method may further including transitioning from the second mode to the first mode when the velocity of the rotorcraft decreases below a second velocity threshold, where transitioning from the second mode to the first mode includes fading in the gain of the first dynamic controller over a third period of time and reactivating the integrator. In some embodiments, the velocity may be a ground speed.

The method may further included fading in a gain of the pilot stick signal in the attitude rate controller over a fourth period of time when a pilot stick of the pilot stick assembly is out of detent; and fading out a gain of the pilot stick signal in the attitude rate controller over a fifth period of time when the pilot stick of the pilot stick assembly is in detent. In some embodiments, the first period of time is a same length as the third period of time, the second period of time is a same length as the fifth period of time, and the first velocity threshold is the same as the second velocity threshold. In some embodiments, the first period of time and the second period of time are between 1 s and 10 s.

In an embodiment, the first dynamic controller further includes a proportional path, and an output of the first dynamic controller is a sum of the proportional path and the output of the integrator. Decreasing the value of the integrator may include coupling the output of the integrator to an input of the integrator via a first feedback path. In some embodiments, the first feedback path includes a limiter.

Another general aspect includes a flight control system for a rotorcraft including: a processor and a non-transitory computer readable storage medium with an executable program stored thereon, where the executable program includes instructions to: receive a pilot control signal via a first interface of the processor; in a first mode determine a first value based on the received pilot control signal, determine a second value based on the determined first value, determine an actuator command based on the determined first value, where determining the actuator command includes executing a first dynamic controller that has an integrator, and providing an output of the first dynamic controller to an actuator via a second interface of the processor, where a state of the rotorcraft corresponding to the first value is configured to be proportional to the received pilot control signal. The method further includes transitioning from the first mode to a second mode when a first condition of the rotorcraft crosses a first predetermined threshold, where transitioning includes fading out a gain of the first dynamic controller over a first period of time, and decreasing a value of the integrator over a second period of time; and in the second mode, providing an output of a second dynamic controller to the actuator via the second interface of the processor, where the output of the second dynamic controller is proportional to the received pilot control signal.

Implementations may include one or more of the following features. The flight control system where the first value is a translational speed, the second value is an attitude, and the first condition of the rotorcraft is a velocity of the rotorcraft. In some embodiments, the first value is a translational speed, the second value is an attitude, and the first condition of the rotorcraft is a velocity of the rotorcraft. The attitude may include a pitch attitude and a roll attitude, executing the first dynamic controller may include executing a first pitch dynamic controller having a pitch integrator and executing a first roll dynamic controller having a roll integrator. Providing the output of the first dynamic controller to the actuator via a second interface of the processor may include providing an output of the first pitch dynamic controller to a pitch actuator and providing an output of the first roll dynamic controller to a roll actuator.

In an embodiment, the executable program further includes instructions to fade in a gain of the pilot control signal in the second dynamic controller when a pilot control stick is out of detent; and fade out a gain of the pilot control signal in the second dynamic controller when the pilot control stick is in detent. The executable program may further include instructions to transition from the second mode to the first mode when the first condition of the rotorcraft decreases below a second predetermined threshold, where transitioning from the second mode to the first mode includes fading in the gain of the first dynamic controller over a third period of time, and reactivating the integrator. In some embodiments, the first predetermined threshold is greater than the second predetermined threshold.

In an embodiment, the first dynamic controller further includes a proportional path, and an output of the first dynamic controller is a sum of the proportional path and the output of the integrator. Decreasing the value of the integrator may include coupling the output of the integrator to an input of the integrator via a first feedback path. In some example, the first feedback path includes a limiter. The first predetermined threshold may include a ground speed and the first period of time and the second period of time may be between 1 s and 10 s.

Another general aspect includes a rotorcraft including: a body; a power train coupled to the body and including a power source and a drive shaft coupled to the power source; a rotor system coupled to the power train and including a plurality of rotor blades; a flight control system operable to change at least one operating condition of the rotor system; a pilot control assembly configured to receive commands from a pilot, where the flight control system is a fly-by-wire flight control system in electrical communication with the pilot control assembly; and a flight control computer in electrical communication between the flight control system and the pilot control assembly. The flight control computer configured to: receive, from the pilot control assembly a pilot command to change a first flight characteristic, when a velocity of the rotorcraft is less than a first velocity threshold, interpret the first flight characteristic as a translational speed in a first mode and determine an attitude based on the translational speed via using an attitude controller including an integrator and a proportional path parallel to the integrator. When the velocity of the rotorcraft is greater than a second velocity threshold, the first flight characteristic is interpreted as a rate in a second mode, and when the velocity of the rotorcraft increases past the second velocity threshold, fade out a gain of the attitude controller is faded out, and a value of the integrator is successively decreased.

Implementations may include one or more of the following features. The rotorcraft where: the attitude includes a pitch attitude; and the rate includes a pitch rate. In the attitude further includes a roll attitude; and the rate further includes a roll rate. The flight control computer may be further configured to transition from the second mode to the first mode when the velocity of the rotorcraft decreases below the first velocity threshold, and transitioning from the second mode to the first mode may include fading in the gain of the attitude controller, and reactivating the integrator. In some embodiments, the first velocity threshold is less than the second velocity threshold.

In an embodiment, the flight control computer is further configured to: fade in a gain of the pilot command to an attitude rate controller when a pilot stick of the pilot control assembly is out of detent; and fade out a gain of the pilot command to the attitude rate controller when the pilot stick of the pilot control assembly is in detent.

Advantages of embodiments include the ability to automatically and smoothly transition between a TRC mode and a rate control mode or between a translational speed hold mode and a rotational rate control mode. Further advantages include making the transition smooth and seamless enough that little change in attitude is noticeable to pilot or occupants of the aircraft.

While this invention has been described with reference to illustrative embodiments, this description is not intended to be construed in a limiting sense. Various modifications and combinations of the illustrative embodiments, as well as other embodiments of the invention, will be apparent to persons skilled in the art upon reference to the description. It is therefore intended that the appended claims encompass any such modifications or embodiments. 

What is claimed is:
 1. A method of operating a rotorcraft, the method comprising: operating the rotorcraft in a first mode comprising determining a translational speed based on a pilot stick signal generated by a pilot stick assembly, determining an attitude based on the determined translational speed, determining an actuator command based on the determined attitude, determining the actuator command comprising using a first dynamic controller having an integrator; providing an output of the first dynamic controller to an actuator, wherein a translational speed of the rotorcraft is proportional to an amplitude of the pilot stick signal in the first mode, and a change in the amplitude of the pilot stick signal causes a change in the translational speed of the rotorcraft; and transitioning from the first mode to a second mode when a velocity of the rotorcraft exceeds a first velocity threshold, transitioning comprising fading out a gain of the first dynamic controller over a first period of time, and decreasing a value of the integrator over a second period of time; and operating the rotorcraft in the second mode comprising providing an output of an attitude rate controller to the actuator, wherein the output of the attitude rate controller is proportional to the amplitude of the pilot stick signal.
 2. The method of claim 1, wherein: determining the translational speed based on the pilot stick signal comprises determining a forward speed component based on cyclic longitude stick control portion of the pilot stick signal and determining a lateral speed component based on a cyclic latitude stick control portion of the pilot stick signal; determining the attitude based on the determined translational speed comprises determining a pitch attitude based on the determined forward speed component and a roll attitude based on the determined lateral speed component; determining the actuator command comprises determining a pitch actuator command based on the determined pitch attitude and a roll actuator command based on the determined roll attitude; the integrator comprises a pitch integrator and a roll integrator; the first dynamic controller comprises a first pitch dynamic controller having the pitch integrator, and a first roll dynamic controller having the roll integrator; providing the output of the first dynamic controller to the actuator comprises providing an output of the first pitch dynamic controller to a pitch actuator and an output of the first roll dynamic controller to a roll actuator; fading out the gain of the first dynamic controller over the first period of time comprises fading out a gain of the first pitch dynamic controller and the first roll dynamic controller over the first period of time; and decreasing the value of the integrator over the second period of time comprises decreasing a value of the pitch integrator and the roll integrator over the second period of time.
 3. The method of claim 2, wherein determining the translational speed further comprises determining a vector sum of the forward speed component and the lateral speed component.
 4. The method of claim 1, wherein: the translational speed comprises a forward speed; the attitude comprises a pitch attitude; the attitude rate controller comprises a pitch rate controller; and the actuator comprises a pitch actuator.
 5. The method of claim 1, further comprising transitioning from the second mode to the first mode when the velocity of the rotorcraft decreases below a second velocity threshold, transitioning from the second mode to the first mode comprising: fading in the gain of the first dynamic controller over a third period of time, and reactivating the integrator.
 6. The method of claim 5, further comprising: fading in a gain of the pilot stick signal in the attitude rate controller over a fourth period of time when a pilot stick of the pilot stick assembly is out of detent; and fading out the gain of the pilot stick signal in the attitude rate controller over a fifth period of time when the pilot stick of the pilot stick assembly is in detent.
 7. The method of claim 6, wherein: the first period of time is a same length as the third period of time; the second period of time is a same length as the fifth period of time; and the first velocity threshold is the same as the second velocity threshold.
 8. The method of claim 5, wherein the first velocity threshold is greater than the second velocity threshold.
 9. The method of claim 1, wherein the first dynamic controller further comprises a proportional path, and an output of the first dynamic controller is a sum of the proportional path and the output of the integrator.
 10. The method of claim 1, wherein decreasing the value of the integrator comprises coupling the output of the integrator to an input of the integrator via a first feedback path.
 11. The method of claim 10, wherein the first feedback path comprises a limiter.
 12. The method of claim 1, wherein the velocity of the rotorcraft comprises a ground speed.
 13. The method of claim 1, wherein the first period of time and the second period of time are between 1 s and 10 s.
 14. A flight control system for a rotorcraft comprising: a processor and a non-transitory computer readable storage medium with an executable program stored thereon, the executable program including instructions to: receive a pilot control signal via a first interface of the processor; in a first mode determine a first value based on the received pilot control signal, determine a second value based on the determined first value, determine an actuator command based on the determined first value, determining the actuator command comprising executing a first dynamic controller that has an integrator, and providing an output of the first dynamic controller to an actuator via a second interface of the processor, wherein a flight state of the rotorcraft corresponding to the first value is configured to be proportional to the received pilot control signal, and a change in the first value is configured to cause a corresponding change in the flight state of the rotorcraft; and transitioning from the first mode to a second mode when a first condition of the rotorcraft crosses a first predetermined threshold, transitioning comprising fading out a gain of the first dynamic controller over a first period of time, and decreasing a value of the integrator over a second period of time; and in the second mode, providing an output of a second dynamic controller to the actuator via the second interface of the processor, wherein the output of the second dynamic controller is proportional to the received pilot control signal.
 15. The flight control system of claim 14, wherein; the first value is a translational speed; the second value is an attitude; and the first condition of the rotorcraft is a velocity of the rotorcraft.
 16. The flight control system of claim 15, wherein: the attitude comprises a pitch attitude and a roll attitude; the integrator comprises a pitch integrator and a roll integrator; executing the first dynamic controller comprises executing a first pitch dynamic controller having the pitch integrator and executing a first roll dynamic controller having the roll integrator; and providing the output of the first dynamic controller to the actuator via the second interface of the processor comprises providing an output of the first pitch dynamic controller to a pitch actuator and providing an output of the first roll dynamic controller to a roll actuator.
 17. The flight control system of claim 14, wherein the executable program further includes instructions to fade in a gain of the pilot control signal in the second dynamic controller when a pilot control stick is out of detent; and fade out the gain of the pilot control signal in the second dynamic controller when the pilot control stick is in detent.
 18. The flight control system of claim 14, wherein the executable program further includes instructions to transition from the second mode to the first mode when the first condition of the rotorcraft decreases below a second predetermined threshold, wherein transitioning from the second mode to the first mode comprises: fading in the gain of the first dynamic controller over a third period of time, and reactivating the integrator.
 19. The flight control system of claim 18, wherein the first predetermined threshold is greater than the second predetermined threshold.
 20. The flight control system of claim 14, wherein the first dynamic controller further comprises a proportional path, and an output of the first dynamic controller is a sum of the proportional path and the output of the integrator.
 21. The flight control system of claim 14, wherein decreasing the value of the integrator comprises coupling the output of the integrator to an input of the integrator via a first feedback path.
 22. The flight control system of claim 21, wherein the first feedback path comprises a limiter.
 23. The flight control system of claim 14, wherein the first predetermined threshold comprises a ground speed.
 24. The flight control system of claim 14, wherein the first period of time and the second period of time are between 1 s and 10 s.
 25. A rotorcraft comprising: a body; a power train coupled to the body and comprising a power source and a drive shaft coupled to the power source; a rotor system coupled to the power train and comprising a plurality of rotor blades; a flight control system operable to change at least one operating condition of the rotor system; a pilot control assembly configured to receive commands from a pilot, wherein the flight control system is a fly-by-wire flight control system in electrical communication with the pilot control assembly; and a flight control computer in electrical communication between the flight control system and the pilot control assembly, the flight control computer configured to: receive, from the pilot control assembly a pilot command to change a first flight characteristic, when a velocity of the rotorcraft is less than a first velocity threshold, interpret the first flight characteristic as a translational speed in a first mode and determine an attitude based on the translational speed via using an attitude controller comprising an integrator and a proportional path parallel to the integrator, when the velocity of the rotorcraft is greater than a second velocity threshold, interpret the first flight characteristic as a rate in a second mode, and when the velocity of the rotorcraft increases past the second velocity threshold, fade out a gain of the attitude controller, and successively decrease a value of the integrator.
 26. The rotorcraft of claim 25, wherein: the attitude comprises a pitch attitude; and the rate comprises a pitch rate.
 27. The rotorcraft of claim 26, wherein: the attitude further comprises a roll attitude; and the rate further comprises a roll rate.
 28. The rotorcraft of claim 25, wherein the flight control computer is further configured to transition from the second mode to the first mode when the velocity of the rotorcraft decreases below the first velocity threshold, and transitioning from the second mode to the first mode comprises: fading in the gain of the attitude controller, and reactivating the integrator.
 29. The rotorcraft of claim 28, wherein the first velocity threshold is less than the second velocity threshold.
 30. The rotorcraft of claim 25, wherein the flight control computer is further configured to: fade in a gain of the pilot command to an attitude rate controller when a pilot stick of the pilot control assembly is out of detent; and fade out the gain of the pilot command to the attitude rate controller when the pilot stick of the pilot control assembly is in detent. 